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Abstract:

A particular method includes exposing a plurality of fibers of a
composite skin of an aircraft component by providing an opening in a
layer of a first resin overlaying the plurality of fibers at an interior
surface of the composite skin. The method also includes bonding a
conductive patch in electrical contact with the plurality of fibers of
the composite skin at the opening using a second resin. The method
further includes preparing the aircraft component to receive a connector
in electrical contact with the conductive patch.

Claims:

1. A method comprising: exposing a plurality of fibers of a composite
skin of an aircraft component by providing an opening in a layer of a
first resin overlaying the plurality of fibers at an interior surface of
the composite skin; bonding a conductive patch in electrical contact with
the plurality of fibers of the composite skin at the opening using a
second resin; and preparing the aircraft component to receive a connector
in electrical contact with the conductive patch.

2. The method of claim 1, further comprising applying a conductive layer
on an external surface of the composite skin.

3. The method of claim 1, wherein a size of the conductive patch is
selected such that a temperature change due to electrical current passing
through the plurality of fibers via the conductive patch is less than
twenty degrees centigrade.

4. The method of claim 1, wherein a size of the conductive patch is
selected such that a temperature change due to electrical current passing
through the plurality of fibers via the conductive patch does not cause a
glass transition temperature of a polymer matrix of the composite skin to
be exceeded.

5. The method of claim 1, wherein preparing the aircraft component to
receive the connector in electrical contact with the conductive patch
includes: forming a hole through the composite skin and the conductive
patch; and inserting a threaded member into the hole, wherein threaded
member is configured to receive the connector.

6. The method of claim 5, wherein forming the hole through the composite
skin and the conductive patch includes drilling the hole and
counterboring the composite skin at an external surface.

7. The method of claim 6, wherein the composite skin includes a
conductive layer at the external surface, and further comprising, after
the threaded member is inserted into the hole, applying a filler putty
between the threaded member and the conductive layer, wherein the filler
putty is substantially flush with the conductive layer.

8. A method comprising: exposing a plurality of fibers of a first section
of a plurality of sections of a composite skin of a fuselage of an
aircraft by providing an opening in a layer of a first resin overlaying
the plurality of fibers at an interior surface of the first section;
bonding a conductive patch in electrical contact with the plurality of
fibers of the first section using a second resin; and preparing the first
section to receive a connector in electrical contact with the conductive
patch.

9. The method of claim 8, further comprising joining the first section
and a second section of the plurality of the sections using a splice
ring, wherein the first section and the second section are adjacent
sections of the fuselage.

10. The method of claim 9, wherein the fuselage includes a conductive
layer at an external surface and wherein the splice ring provides
electrical contact between a portion of the conductive layer of the first
section and a portion of the conductive layer of the second section.

11. The method of claim 9, wherein the splice ring provides electrical
contact between the plurality of fibers of the first section and a
plurality of fibers of the second section.

12. The method of claim 11, further comprising: applying a first bonding
plate in electrical contact with the plurality of fibers of the first
section; and applying a second bonding plate in electrical contact with
the plurality of fibers of the second section, wherein the splice ring
electrically couples the first bonding plate and the second bonding
plate.

13. The method of claim 8, further comprising: coupling the connector in
electrical contact with the conductive patch; and coupling a powered
device of the aircraft to the connector via a first conductor; wherein
the powered device is configured to receive electrical current from a
power source of the aircraft via a second conductor and to return
electrical current to the power source via the first conductor and the
plurality of fibers.

14. The method of claim 13, wherein an external surface of the fuselage
includes a conductive layer, and wherein the electrical current from the
powered device passes through the plurality of fibers to the conductive
layer.

15. A method comprising: coupling a powered device to a power source via
a conductor, wherein the power source provides electrical current to the
powered device via the conductor; and coupling the powered device to the
power source through a composite structure, wherein the power source
receives return current from the powered device via the composite
structure, wherein the composite structure includes a first resin layer
overlaying a plurality of fibers, and wherein the powered device is
coupled to the composite structure via a conductive patch that is bonded
in electrical contact with the plurality of fibers by a second resin at
an opening in the first resin layer.

16. The method of claim 15, wherein the composite structure includes a
first surface and a second surface, wherein the patch is coupled at least
partially within the opening on the first surface and a conductive layer
is coupled to the second surface.

17. The method of claim 16, wherein the return current from the powered
device passes from the conductive patch through the plurality of fibers
to the conductive layer.

18. The method of claim 15, further comprising coupling the powered
source to the composite structure via a second conductive patch that is
in electrical contact with the plurality of fibers at a second opening in
the first resin layer.

19. The method of claim 15, wherein the plurality of fibers includes
boron fibers, carbon fibers, tungsten fibers, or a combination thereof.

20. The method of claim 15, wherein the conductive patch is an expanded
metal patch.

Description:

CLAIM OF PRIORITY

[0001] This application is a continuation patent application of, and
claims priority from U.S. patent application Ser. No. 12/393,695, filed
on Feb. 26, 2009 and entitled "DISTRIBUTING POWER IN SYSTEMS HAVING A
COMPOSITE STRUCTURE" which is incorporated by reference herein in its
entirety for all purposes.

FIELD OF THE DISCLOSURE

[0002] The present disclosure is generally related to composite structure
power distribution.

BACKGROUND

[0003] Power distribution systems can add significant weight and
complexity to any system. For example, the power distribution system of a
large commercial aircraft may be a significant weight burden for the
aircraft. Power distribution systems typically use at least two wires to
supply power from a power source to a powered device: a first wire (or a
powered wire) to send current from the power source to the powered device
and a second wire (or a return wire) to receive return current from the
powered device at the power source. Thus, the power distribution system
includes the weight of both the power wire and the return wire. Besides
the issue of the weight of the two or more wires, using two or more wires
to distribute power has the additional concern of routing the wires. In
an aircraft or other vehicle where conductive metals are the primary
structure, a one-wire power distribution system can be used. For example,
current may be sent from the power source to the powered device and
returned to the power source via the conductive metal primary structure.
In any case, for aircraft and other vehicles, wire routing can be a
significant concern due to issues such as shielding the wires to avoid
chafing and routing the wires in safe locations.

SUMMARY

[0004] Systems and methods to distribute power in systems having a
composite structure are disclosed. In a particular embodiment, a method
of distributing power in a system having a composite structure includes
coupling a powered device to a power source via a conductor. The power
source provides electrical current to the powered device via the
conductor. The method also includes coupling the powered device to the
power source through a composite structure via a patch that is in direct
contact with fibers of the composite structure. The power source receives
return current from the powered device via the composite structure.

[0005] In another particular embodiment, an aircraft includes a fuselage
having a composite skin. The composite skin includes a first resin and a
plurality of fibers. The fuselage also includes an expanded conductive
patch bonded to an interior surface of the composite skin by a second
resin. The expanded conductive patch is in electrical contact with the
plurality of fibers. The fuselage also includes a connector in electrical
contact with the expanded conductive patch.

[0006] In another particular embodiment, a system includes at least one
composite layer comprising a first resin and a plurality of fibers. The
system also includes an expanded conductive patch bonded to a first
surface of the at least one composite layer by a second resin. The
expanded conductive patch is in electrical contact with the plurality of
fibers. The system further includes a connector in electrical contact
with the expanded conductive patch.

BRIEF DESCRIPTION OF THE DRAWINGS

[0007]FIG. 1 is a flow diagram of an aircraft production and service
methodology;

[0008]FIG. 2 is a block diagram of functional groups of aircraft systems;

[0009] FIG. 3 is an illustration of a first embodiment of a system to
distribute power;

[0010] FIG. 4 is an illustration of a second embodiment of a system to
distribute power;

[0011] FIG. 5 is an illustration of a third embodiment of a system to
distribute power;

[0012]FIG. 6 is an illustration of an aircraft having a composite
structure;

[0013]FIG. 7 is flow diagram of a particular embodiment of a method of
distributing power in a system having a composite structure; and

[0014]FIG. 8 is a flow diagram of a particular embodiment of a method of
making a power distribution system.

DETAILED DESCRIPTION

[0015] The features, functions, and advantages that are discussed can be
achieved independently in various embodiments disclosed herein or may be
combined in yet other embodiments further details of which can be shown
with reference to the following description and drawings.

[0016] Power distribution systems that enable routing return current via a
composite structure are described. For example, a composite structural
member, such as a composite skin of an aircraft, may be used to send
return current from a powered device to a power source. Use of such power
distribution systems may lead to significant weight reductions since no
return wire or network of conductors is required. Such power distribution
systems may also use less space than two wire systems (i.e., systems that
use a return wire). Additionally, these power distribution systems may be
less complex to install than power distribution systems that use a return
wire because the return wire does not need to be routed or otherwise
protected against chafing or other potential destructive influences.

[0017] Referring more particularly to the drawings, embodiments of the
disclosure may be described in the context of an aircraft manufacturing
and service method 100 as shown in FIG. 1 and functional aspects of an
aircraft 200 as shown in FIG. 2. During pre-production, an exemplary
method 100 may include specification and design 110 of the aircraft 200
and material procurement 120. During production, component and
subassembly manufacturing 130 and system integration 140 of the aircraft
200 takes place. Thereafter, the aircraft 200 may go through
certification and delivery 150 in order to be placed in service 160.
While in service by a customer, the aircraft 200 is scheduled for routine
maintenance and service 170 (which may also include modification,
reconfiguration, refurbishment, and so on).

[0018] Each of the processes of the method 100 may be performed or carried
out by a system integrator, a third party, and/or an operator (e.g., a
customer). For the purposes of this description, a system integrator may
include without limitation any number of aircraft manufacturers and
major-system subcontractors; a third party may include without limitation
any number of vendors, subcontractors, and suppliers; and an operator may
be an airline, a leasing company, a military entity, a service
organization, and so on.

[0019] As shown in FIG. 2, the aircraft 200 produced by the exemplary
method 100 may include an airframe 210 with a plurality of systems 220
and an interior 230. Examples of high-level systems 220 include one or
more of a propulsion system 240, an electrical system 250, a hydraulic
system 260, and an environmental system 270. Any number of other systems
may be included. Although an aerospace example is shown, the principles
of the various embodiments may be applied to other industries, such as
the automotive industry, the ship building industry or the wind power
industry.

[0020] Apparatus and methods embodied herein may be employed during any
one or more of the stages of the production and service method 100. For
example, components or subassemblies corresponding to the component and
subassembly manufacturing process 130 may be fabricated or manufactured
in a manner similar to components or subassemblies produced while the
aircraft 200 is in its in service phase 160. Also, one or more apparatus
embodiments, method embodiments, or a combination thereof may be utilized
during the production stages, such as the component and subassembly
manufacturing stage 130 or the system integration stage 140, for example,
by substantially expediting assembly of or reducing the cost of the
aircraft 200. Similarly, one or more of the apparatus embodiments, the
method embodiments, or a combination thereof may be utilized while the
aircraft 200 is in service, for example and without limitation, during
the maintenance and service stage 170.

[0021] FIG. 3 is an illustration of a first embodiment of a system to
distribute power. The system includes a powered device 306 coupled via a
first conductor 304 to a power source 302. The powered device 306 is also
coupled to a connector 322 supported by a conductive patch 320 on a
composite member 310 via a second conductor 308. In a particular
embodiment, the composite member 310 may be a panel or other member of an
aircraft, such as the aircraft 200 of FIG. 2. In this embodiment, the
powered device 306 may be an aircraft component that receives power from
the power source 302 via the first conductor 304.

[0022] In a particular embodiment, the composite member 310 includes a
plurality of fibers bound by one or more resins. For example, the
composite member 310 may include a carbon fiber reinforced polymer (CFRP)
member, a composite of non-carbon fibers (e.g., boron fibers, metal
fibers, etc.) bound by a resin or other continuous medium. For purposes
of illustration in FIG. 3, the composite member 310 is illustrated having
a resin layer 312 and a plurality of fiber plies, including a first fiber
ply 314 and a second fiber ply 316. However, the composite member 310 may
include more than one resin, more than two fiber plies, fibers arranged
in another manner, such as individual fibers, fiber mats, fiber cloth,
and so forth. Additionally, while the resin layer 312 is shown in FIG. 3
at the surface of the composite member 310, a resin may bind the
plurality of fibers throughout the composite member 310. For example, the
resin may be a substantially continuous medium in which the fiber plies
314, 316 reside.

[0023] In a particular embodiment, the composite member 310 also has a
conductive layer 318 applied to a surface of the composite member 310.
For example, where the composite member 310 is a portion of the aircraft,
e.g., a portion of an aircraft panel or structural member, the conductive
layer 318 may include a lightning protection layer. In a particular
illustrative embodiment, the conductive layer 318 may include a metal
layer or a metal foil that substantially covers an entire surface of the
composite member 310.

[0024] In a particular embodiment, the powered device 306 is coupled to
the composite member 310 via the connector 322 of the conductive patch
320. In a particular embodiment, the conductive patch 320 includes an
expanded conductive patch. Expanded, in this context, refers to having
gaps or openings between portions of the patch. For example, an expanded
conductive patch may include a mesh, a grid, or a fabric of conductive
elements, a member having a plurality of holes or openings, another
structure where individual conductive elements or portions of conductive
elements are spaced apart, or any combination thereof. The openings
enable the expanded conductive patch to be coupled to fibers of the first
fiber ply 314 using a resin. To illustrate, the openings enable the resin
to flow around or between the individual conductive elements (e.g., metal
strands). Thus, when put under compaction the individual conductive
elements may come into direct and intimate contact with the fibers of the
structure immediately below (such as the first fiber ply 314). If a solid
foil is used, the resin may pool beneath the foil and inhibit the foil
from coming into direct and intimate contact with the fibers.

[0025] In an alternative embodiment, the conductive patch 320 is
corrugated, channelized, or otherwise structured to allow the resin to
flow out of the space between the fibers of the first fiber ply 314 and
the conductive patch 320 to enable direct and intimate contact between
the fibers and the conductive patch 320. To illustrate, the conductive
patch 320 can be placed in direct contact with the fibers of the first
fiber ply 314 and the resin can be applied over the conductive patch 320
to adhere the conductive patch 320 to the first fiber ply 314. In a
particular embodiment, electrical current may flow from the conductive
patch 320 to the conductive layer 318 via the fiber plies 314, 316.
Additionally, some of the return current may flow along a length of the
fibers of the fiber plies 314, 316 to the power source 302. In a
particular embodiment, the power source 302 may be coupled via a second
conductive patch (not shown) to the composite member 310 or to another
composite member (not shown) near the power source 302. The composite
member near the power source 302 may be in electrical contact with the
conductive layer 318, the fiber plies 314, 316, or both. Accordingly,
return current flow to the power source 302 is provided via the
conductive patch 320.

[0026] In a particular embodiment, the conductive patch 320 is deposited
directly onto the exposed fibers of the first fiber ply 314. For example,
a metallic patch may be deposited onto the fibers by first exposing the
fibers by chemical or physical removal of the resin and then applying a
metal layer to the fibers. The metal layer may be deposited using
chemical, physical or electrical processes, such as electrodepositing or
electroplating, liquid or gaseous chemical deposition of metals,
sputtering, plasma or arc spraying metal onto the fibers, or the like.

[0027] The second conductor 308 may be coupled to the expanded conductive
patch 320 via the connector 322. For example, the connector 322 may
include a threaded member physically coupled to the composite member 310
and in electrical contact with the expanded conductive patch 320. In a
particular embodiment, the expanded conductive patch 320 enables return
current to flow from the powered device 306 to the power source 302 via
the fiber plies 314, 316. For example, electrical current supplied to the
powered device 306 via the first conductor 304 may be returned to the
power source 302 via the expanded conductive patch 320. The expanded
conductive patch 320 may be in physical contact with the fibers of at
least the first fiber ply 314.

[0028] FIG. 4 is an illustration of a second embodiment of a system to
distribute power. The system illustrated in FIG. 4 includes a composite
member 400. In a particular embodiment, the composite member 400 includes
a skin or structural member of a vehicle having a composite construction.
For example, the composite member 400 may include a panel, a structural
member, or a fuselage segment of an aircraft, such as the aircraft 200 of
FIG. 2.

[0029] The composite member 400 may include a plurality of fibers, such as
a first fiber ply 404 and a second fiber ply 406, bound by a polymer
matrix. In FIG. 4, the polymer matrix is illustrated as a resin layer
402; however, the polymer matrix may be at a surface or surfaces of the
composite member, intermixed with the plurality of fibers, at other
locations of the composite member 400, or any combination thereof. In a
particular embodiment, the polymer matrix is a continuous medium and the
fibers of the fiber plies 404, 406 are a discontinuous medium of the
composite member 400. In a particular embodiment, the composite member
400 includes a conductive layer 408. For example, the conductive layer
408 may include a lightning protection layer of an aircraft. To
illustrate, the composite member 400 may include an internal surface 410
on the inside of the aircraft and an external surface 412 on the outside
of the aircraft. The external surface 412 may be covered by a
substantially continuous metal conductive layer.

[0030] In a particular embodiment, the composite member 400 includes a
conductive patch 420 at the internal surface 410. The conductive patch
420 may be electrically bonded to (e.g., in physical contact with) fibers
of at least one of the fiber plies 404, 406. For example, the conductive
patch 420 may rest on and be in physical contact with at least the first
fiber ply 404. Additionally, the conductive patch 420 may be in
electrical contact with a connector 426. The connector 426 may include a
washer 424 to increase the electrical contact area with the conductive
patch 420.

[0031] In a particular embodiment, the connector 426 may be coupled to the
composite member 400 via a threaded insert 450. The threaded insert 450
may be inserted into a counterbore 452 through the external surface 412
of the composite member 400. In a particular embodiment, where it is
desirable for the external surface 412 to be smooth, such as for the
external surface of an aircraft, the counterbore 452 may be filled with a
filler putty 454. The filler putty 454 may fill a gap between the
counterbore 452 and the conductive layer 408 and may be substantially
flush with the external surface 412. In this embodiment, the threaded
insert 450 and the connector 426 may not have a direct electrical
communication path to the conductive layer 408.

[0032] The conductive patch 420 may be coupled to the composite member 400
via a resin 422. The resin 422 may be applied over the conductive patch
420 to bond the conductive patch 420 to the plurality of fibers of the
first fiber ply 404. In a particular embodiment, the conductive patch 420
includes a plurality of holes to allow the resin 422 to contact the first
fiber ply 404 while the conductive patch 420 is also in electrical
contact with the first fiber ply 404. In a particular embodiment, the
conductive patch 420 includes an expanded metal patch. Expanded, in this
context, indicates that portions of the conductive patch 420 are
separated from one another in a manner that allows the resin 422 to
couple to the first fiber ply 404 while the conductive patch 420 is also
in contact with the first fiber ply 404. The expanded metal patch may
include a metal mesh, metallic cloth, metal foil with a plurality of
holes, or other structure that provides holes for the resin 422 to pass
through to contact the plurality of fibers of the first fiber ply 404
when the conductive patch 420 is in direct contact with the plurality of
fibers of the first fiber ply 404.

[0033] In a particular embodiment, the first fiber ply 404 and the second
fiber ply 406 include boron fibers, carbon fibers, tungsten fibers or
other fibers. The conductive patch 420 may be formed of a material that
is galvanically compatible with the plurality of fibers of the fiber
plies 404, 406, with the resin layer 402 and with the resin 422. For
example, the conductive patch 420 may include copper, nickel, steel,
another galvanically compatible metal or alloy, or any combination
thereof.

[0034] A size of the conductive patch 420 may be selected based on an
amount of current to be routed through the conductive patch 420. For
example, the size of the conductive patch 420 may be related to a current
density of the return current passing through the conductive patch 420 to
the fibers of the fiber plies 404, 406. The current density may also be
related to a temperature rise that occurs when the return current passes
through conductive patch 420 to the fibers of the fiber plies 404, 406.
In a particular embodiment, the size of the conductive patch 420 may be
selected such that, during operation, a temperature change due to the
return electrical current passing through the plurality of fibers of the
first fiber ply 404 and the second fiber ply 406 via the conductive patch
420 does not cause a glass transition temperature of the resin layer 402
or of the resin 422 to be exceeded. For example, the size of the
conductive patch 420 may be selected such that, during operation, a
temperature change due to the return current passing through the
plurality of fibers of the first fiber ply 404 and the second fiber ply
406 from the conductive patch 420 is less than about 20 degrees
centigrade.

[0035] In a particular embodiment, return current received via the
connector 426 from a powered device via a wire 470 is passed through the
conductive patch 420 to the plurality of fibers of the first fiber ply
404 and the second fiber ply 406 to the conductive layer 408. This
arrangement allows the powered device to pass return current to a power
source via the conductive layer 408 and through the plurality of fibers
of the first fiber ply 404 and the second fiber ply 406 without requiring
a conductor separate from the composite member 400 to return current from
the powered device to the power source. In an illustrative embodiment,
the return current flows through a z-axis 460 direction of the fiber
plies 404, 406. The z-axis 460 direction refers to the direction from the
conductive patch 420 to the conductive layer 408 (e.g., through a
thickness of the composite member 400.

[0036] FIG. 5 is an illustration of a second embodiment of a system to
distribute power. The system illustrated in FIG. 5 includes a composite
member 500. In a particular embodiment, the composite member 500 includes
a skin or structural member of a vehicle having a composite construction.
For example, the composite member 500 may include a panel, a structural
member, or a fuselage segment of an aircraft, such as the aircraft 200 of
FIG. 2.

[0037] The composite member 500 may include a plurality of fibers, such as
a first fiber ply 504 and a second fiber ply 506, bound by a polymer
matrix. In FIG. 5, the polymer matrix is illustrated as a resin layer
502; however, the polymer matrix may be at a surface or surfaces of the
composite member 500, intermixed with the plurality of fibers, at other
locations of the composite member 500, or any combination thereof. In a
particular embodiment, the polymer matrix is a continuous medium and the
fibers of the fiber plies 504, 506 are a discontinuous medium of the
composite member 500. In a particular embodiment, the composite member
500 includes a conductive layer 508. For example, the conductive layer
508 may include a lightning protection layer of an aircraft. To
illustrate, the composite member 500 may include an internal surface 510
on the inside of the aircraft and an external surface 512 on the outside
of the aircraft. The external surface 512 may be covered by a metal
conductive layer.

[0038] In a particular embodiment, the composite member 500 includes a
conductive patch 520 at the internal surface 510. The conductive patch
520 may be electrically bonded to (e.g., in physical contact with) fibers
of at least one of the fiber plies 504, 506. For example, the conductive
patch 520 may rest on and be in physical contact with at least the first
fiber ply 504. Additionally, the conductive patch 520 may be in
electrical contact with a flush connector 526. The flush connector 526
may include a hi-lock type fastener with a ring terminal 524 on the
inside. The flush connector 526 maybe attached to the composite member
500 via a countersink 552 in the external surface 512 such that the flush
connector 526 is substantially flush with the external surface 512.

[0039] The conductive patch 520 may be coupled to the composite member 500
via a resin 522. The resin 522 may be applied over the conductive patch
520 to bond the conductive patch 520 to the plurality of fibers of the
first fiber ply 504. In a particular embodiment, the conductive patch 520
includes a plurality of holes to allow the resin 522 to contact the first
fiber ply 504 while the conductive patch 520 is also in electrical
contact with the first fiber ply 504. In a particular embodiment, the
conductive patch 520 includes an expanded metal patch. Expanded, in this
context, indicates that portions of the conductive patch 520 are
separated from one another in a manner that allows the resin 522 to
couple to the first fiber ply 504 while the conductive patch 520 is also
in contact with the first fiber ply 504. The expanded metal patch may
include a metal mesh, metallic cloth, metal foil with a plurality of
holes, or other structure that provides holes for the resin 522 to pass
through to contact the plurality of fibers of the first fiber ply 504
when the conductive patch 520 is in direct contact with the plurality of
fibers of the first fiber ply 504.

[0040] In a particular embodiment, the first fiber ply 504 and the second
fiber ply 506 include boron fibers, carbon fibers, tungsten fibers or
other fibers. The conductive patch 520 may be formed of a material that
is galvanically compatible with the plurality of fibers of the fiber
plies 504, 506, with the resin layer 502 and with the resin 522. For
example, the conductive patch 520 may include copper, nickel, steel,
another galvanically compatible metal or alloy, or any combination
thereof.

[0041] A size of the conductive patch 520 may be selected based on an
amount of current to be routed through the conductive patch 520. For
example, the size of the conductive patch 520 may be related to a current
density of the return current passing through the conductive patch 520 to
the fibers of the fiber plies 504, 506. The current density may also be
related to a temperature rise that occurs when the return current passes
through conductive patch 520 to the fibers of the fiber plies 504, 506.
In a particular embodiment, the size of the conductive patch 520 may be
selected such that, during operation, a temperature change due to the
return electrical current passing through the plurality of fibers of the
first fiber ply 504 and the second fiber ply 506 via the conductive patch
520 does not cause a glass transition temperature of the resin layer 502
or of the resin 522 to be exceeded. For example, the size of the
conductive patch 520 may be selected such that, during operation, a
temperature change due to the return current passing through the
plurality of fibers of the first fiber ply 504 and the second fiber ply
506 from the conductive patch 520 is less than about 20 degrees
centigrade.

[0042] The flush connector 526 may be used to couple a powered device to
the conductive patch 520. For example, a threaded portion 550 of the
flush connector 526 may receive the ring terminal 524, a nut 572, another
device to couple a wire to the powered device 570 to the flush connector
526, or any combination thereof. In a particular embodiment, return
current received via the wire 570 from the powered device is passed
through the conductive patch 520 to the plurality of fibers of the first
fiber ply 504 and the second fiber ply 506 to the conductive layer 508.
This arrangement allows the powered device to pass return current to a
power source via the conductive layer 508 and through the plurality of
fibers of the first fiber ply 504 and the second fiber ply 506 without
requiring a conductor separate from the composite member 500 to return
current from the powered device to the power source. In an illustrative
embodiment, the return current flows through a z-axis 560 direction of
the fiber plies 504, 506. The z-axis 560 direction refers to the
direction from the conductive patch 520 to the conductive layer 508
(e.g., through a thickness of the composite member 500.

[0043]FIG. 6 is an illustration of an aircraft 600 having a composite
structure. The aircraft 600 is shown partially deconstructed to highlight
various components. The aircraft 600 includes a fuselage 602 that
includes a plurality of sections, such as a first section 604 and a
second section 606. The first section 604 and the second section 606
include composite members. The composite members may be similar to the
composite members 310, 400 and 500, discussed with reference to FIG. 3,
FIG. 4 and FIG. 5, respectively. For example, the composite member may
include a plurality of fibers 642 that are bound together by a resin (not
shown in FIG. 6).

[0044] In a particular embodiment, the aircraft 600 includes a power
source 622 coupled via a conductor 624 (such as a wire) to a powered
device 620. The power source 622 and the powered device 620 may be in the
same section 604, 606, or in different sections 604, 606. The powered
device 620 may be coupled via a first patch 640 to the plurality of
fibers 642. In a particular embodiment, the power source 622 supplies
current to the powered device 620 via the conductor 624 and the powered
device 620 provides return current to the power source 622 via the
plurality of fibers 642. For example, the powered device 620 may send the
return current via the first patch 640 to the plurality of fibers 642.
The power source 622 may be coupled to the plurality of fibers 642 via a
second patch 660.

[0045] In a particular embodiment, the plurality of fibers 642 of the
second section 606 may be coupled to fibers (not shown) of the first
section 604 via one or more bonding plates. For example, the first
section 604 may include a first bonding plate 644. The first bonding
plate 644 may include a conductive patch bonded to the fibers of the
first section 604. That is, the first bonding plate 644 may be in
electrical contact with the plurality of fibers of the first section 604.
Additionally, the second section 606 may include a second bonding plate
646. The second bonding plate 646 may include a conductive patch in
electrical contact with the plurality of fibers 642 of the second section
606. A splice ring 648 may electrically couple the first bonding plate
644 to the second bonding plate 646 creating an electrical union between
the second section 606 and the first section 604 to enable current to
flow from the plurality of fibers 642 of the second section 606 to the
fibers of the first section 604.

[0046] In a particular embodiment, the first section 604, the second
section 606 or other sections of the aircraft 600 may include a lightning
protection layer on an outer surface. For example, the lightning
protection layer may include a conductive layer, such as the conductive
layer 318 described with reference to FIG. 3, the conductive layer 408
described with reference to FIG. 4, or the conductive layer 508 described
with reference to FIG. 5. The splice ring 648 may provide an electrical
union between the lightning protection layer of the first section 604 and
the lightning protection layer of the second section 606 or lightning
protection layers of other sections of the aircraft 600. In FIG. 6, the
aircraft 600 has been illustrated as having two sections (e.g., the first
section 604 and the second section 606); however, the aircraft 600 may
include any number of sections. Any two or more of the sections of the
aircraft 600 may be electrically coupled via bonding plates and splice
rings as illustrated with respect to the first section 604 and the second
section 606 to provide electrical contact between fibers of each section,
to provide electrical contact between lightning protection layers
associated with each section, or both. Thus, current may flow from the
power source 622 to powered devices, such as the powered device 620, at
any location in the aircraft 600 using a powered wire, such as the
conductor 624, without a return conductor. The return current may be
provided from the powered device 620 to the power source 622 via the
plurality of fibers of the aircraft 600. In a particular embodiment, the
return current may pass through the plurality of fibers to the lightning
protection layer, and the lightning protection layer may provide an
electrical path to communicate the return current to the power source
622.

[0047] In a particular embodiment, a size of each of the conductive
patches, such as the first patch 640, the second patch 660, the bonding
plates 644, 646, or any combination thereof may be selected based on
materials used in the aircraft 600. For example, the sizes of the
conductive patches 640, 660 and the bonding plates 644, 646 may be
selected based on a resin used to bind the fibers 642 in the composite
members. In another example, the sizes of the conductive patches 640, 660
and the bonding plates 644, 646 may be selected based on a resin used to
bind the conductive patches 640, 660 and the bonding plates 644, 646 to
the fibers 642. In an illustrative embodiment, the sizes of the
conductive patches 640, 660 and the bonding plates 644, 646 are selected
such that a temperature change due to current passing through the fibers
via the conductive patches 640, 660 and the bonding plates 644, 646 does
not cause a glass transition temperature of a resin to be exceeded. For
example, the size of each of the conductive patches 640, 660 and the size
of each of the bonding plates 644, 646 may be selected such that a
temperature change due to the return current passing through the fibers
via the conductive patches 640, 660, and the bonding plates 644, 646 is
less than about 20 degrees centigrade.

[0048]FIG. 7 is flow diagram of a particular embodiment of a method of
distributing power in a system having a composite structure. The method
described with reference to FIG. 7 may be used in conjunction with a
composite member, such as the composite member 310 described with
reference to FIG. 3, the composite member 400 described with reference to
FIG. 4, or the composite member 500 described with reference to FIG. 5.
Further, the method may be used in conjunction with a system having a
composite structure, such as the aircraft 600 discussed with reference to
FIG. 6.

[0049] The method includes, at 702, coupling a powered device to a power
source via a conductor. The power source provides electrical current to
the powered device via the conductor. The method includes, at 704,
coupling the powered device to the power source through the composite
structure via a patch that is in direct contact with fibers of the
composite structure. The power source receives return current from the
powered device via the composite structure. In a particular embodiment,
the composite structure includes a first surface and a second surface.
The patch may be coupled to the first surface and a substantially
continuous conductive layer may be coupled to the second surface. For
example, the first surface may include an interior portion of an aircraft
and the second surface may include an exterior portion of the aircraft.
To illustrate, the substantially continuous conductive layer may include
a lightning protection layer. In a particular embodiment, the return
current flows through the fibers of the composite structure to the
substantially continuous conductive layer from the powered device to the
power source. In another particular embodiment, the return current flows
through the fibers of the composite structure from the powered device to
the power source.

[0050]FIG. 8 is a flow diagram of a particular embodiment of a method of
making a power distribution system. The method described with reference
to FIG. 8 may be used in conjunction with a composite member, such as the
composite member 310 described with reference to FIG. 3, the composite
member 400 described with reference to FIG. 4, or the composite member
500 described with reference to FIG. 5. Further, the method may be used
in conjunction with a system having a composite structure, such as the
aircraft 600 described with reference to FIG. 6. Additionally, a power
distribution system made according to the method described with reference
to FIG. 8 may be used to distribute power as discussed with reference to
FIG. 7.

[0051] The method includes, at 802, exposing a plurality of fibers of a
composite member. For example, the composite member may include a
plurality of fibers bound by a resin. In this example, the plurality of
fibers may be covered by the resin and exposing the plurality of fibers
may include grinding away or otherwise removing a portion of the resin to
enable direct access to the plurality of fibers.

[0052] The method also includes, at 804, applying an expanded conductive
patch to the plurality of fibers of the composite member. For example,
after exposing the plurality of fibers, the expanded conductive patch may
be placed in direct contact with the fibers. The expanded conductive
patch may include a conductor that is corrosion resistant and is selected
to be galvanically compatible with the composite member. For example, the
expanded conductive patch may include copper, nickel, steel, another
galvanically compatible metal or alloy, or any combination thereof.

[0053] At 806, a resin may be applied to the expanded conductive patch to
bond the expanded conductive patch in direct contact with the plurality
of fibers. The expanded conductive patch may also be selected to be
galvanically compatible with the resin. The resin may be cured, at 808.
For example, the composite member may be placed in an autoclave at a
specified temperature and pressure to cure the resin. In another example,
the expanded conductive patch and the resin may be pressed onto the
composite member using a vacuum bag and heated to cure the resin and bond
the expanded conductive patch onto the composite member. To illustrate,
such a vacuum bag and heating system may be used to apply the expanded
conductive patch to an aircraft in service.

[0054] At 810, a portion of the resin bonding the expanded conductive
patch to the composite member may be removed to expose a portion of the
expanded conductive patch. For example, a portion of the resin may be
ground away, chemically stripped, peeled away or otherwise removed. At
812, the composite member may be drilled and counterbored on a second
side of the composite member, where the second side is opposite a side on
which the expanded conductive patch is applied. A threaded member may be
inserted into the composite member via the counterbore, at 814. An
electrical connector may be coupled to the threaded member through the
composite member, at 816. Additionally, in a particular embodiment, a
filler putty may be provided to fill the counterbore to smooth the second
surface of the composite member.

[0055] The conductive patch may provide an electrical path from the
electrical connector to the plurality of fibers of the composite member.
In a particular embodiment, the electrical connector may be coupled to a
powered device to provide return current to a power source via the
composite member. For example, the composite member may be a portion of a
panel of an aircraft, and the plurality of fibers may be used to provide
a return current path to the power source from the powered device via the
conductive patch.

[0056] A size of the expanded conductive patch may be selected to reduce
peak current density by spreading out the return current over an area of
the expanded conductive patch. For example, if only a bolt were used to
electrically couple the powered device to the fibers, a smaller contact
area between the fibers and the bolt would be achieved than the contact
area between the expanded conductive patch and the fibers. The smaller
contact area results in a higher current density. The current density
associated with the bolt may be high enough that a temperature change
that results from passing current from the bolt to the fibers becomes
undesirable. For example, the temperature may approach a glass transition
temperature associated with the composite member. Approaching or
exceeding the glass transition temperature may cause the composite member
to soften and would be undesirable. By spreading the current out over the
expanded conductive patch, the local current density can be reduced,
thereby reducing the temperature change associated with the current. To
illustrate, an operational temperature of a carbon fiber reinforced
polymer resin may be about 72 degrees centigrade. To ensure that the
glass transition temperature is not reached, an upper safe temperature
rise limit may be set at about 20 degrees centigrade. Accordingly, the
expanded conductive patch may be sized such that the temperature change
resulting from passing current through the expanded conductive patch is
less than about 20 degrees centigrade. Thus, the expanded conductive
patch can be used to safely distribute the current over a wide area.
Additionally, when a conductive layer is present, the current can be
passed through the fibers of the composite member to the conductive layer
without undesirable temperature increases.

[0057] In a particular embodiment, a similar arrangement to that described
above may be used to couple a power source to the composite member or
other composite members of the system. Since the power source may be
receiving return current from a number of powered devices, an expanded
conductive patch used to couple the power source to the composite member
may be larger than the expanded conductive patch used to couple each of
the powered devices to the composite member. For example, when the
composite member is part of an aircraft, the expanded conductive patch
used to receive the return current at the power source may include a ring
around a section of the fuselage.

[0058] In a particular embodiment, the return current can be passed to the
lightning protection layer of an aircraft so that the return current can
flow back to the power source. Using the expanded conductive patch
described herein allows the current to flow to the lightning protection
layer in a manner that has an acceptable current density and does not
increase drag of the aircraft. For example, the expanded conductive patch
can be coupled to an inner surface of the aircraft. Additionally, a
coupler to connect a powered device to the expanded conductive patch can
be attached to a threaded insert in a counterbore through an exterior
surface of the aircraft. Further, a filler putty may be used to cover the
counterbore. Thus, the threaded insert does not increase drag
significantly, if at all. Because current is passed from the expanded
conductive patch to the lightning protection layer, there is no need to
precisely install the threaded insert to achieve electrical contact
between the threaded insert and the lightning protection layer.

[0059] Although the systems and methods to distribute power disclosed
above have primarily been discussed in the context of use in an aircraft,
the systems and methods are also useful in other areas. For example, the
disclosed systems and methods to distribute power may be useful in an
environment where power is to be distributed in a system that includes a
composite structure. Examples of such systems include, but are not
limited to: windmills and other composite power generating devices;
automobiles, boats and other vehicles.

[0060] The illustrations of the embodiments described herein are intended
to provide a general understanding of the structure of the various
embodiments. The illustrations are not intended to serve as a complete
description of all of the elements and features of apparatus and systems
that utilize the structures or methods described herein. Many other
embodiments may be apparent to those of skill in the art upon reviewing
the disclosure. Other embodiments may be utilized and derived from the
disclosure, such that structural and logical substitutions and changes
may be made without departing from the scope of the disclosure.
Additionally, the illustrations are merely representational and may not
be drawn to scale. Certain proportions within the illustrations may be
exaggerated, while other proportions may be reduced. Accordingly, the
disclosure and the figures are to be regarded as illustrative rather than
restrictive.

[0061] Although specific embodiments have been illustrated and described
herein, it should be appreciated that any subsequent arrangement designed
to achieve the same or similar purpose may be substituted for the
specific embodiments shown. This disclosure is intended to cover any and
all subsequent adaptations or variations of various embodiments.
Combinations of the above embodiments, and other embodiments not
specifically described herein, will be apparent to those of skill in the
art upon reviewing the description.

[0062] The Abstract of the Disclosure is provided with the understanding
that it will not be used to interpret or limit the scope or meaning of
the claims. In addition, in the foregoing Detailed Description, various
features may be grouped together or described in a single embodiment for
the purpose of streamlining the disclosure. This disclosure is not to be
interpreted as reflecting an intention that the claimed embodiments
require more features than are expressly recited in each claim. Rather,
as the following claims reflect, claimed subject matter may be directed
to less than all of the features of any of the disclosed embodiments.
Thus, the following claims are incorporated into the Detailed
Description, with each claim standing on its own as defining separately
claimed subject matter.

[0063] The above-disclosed subject matter is to be considered
illustrative, and not restrictive, and the appended claims are intended
to cover all such modifications, enhancements, and other embodiments,
which fall within the scope of the present disclosure. Thus, to the
maximum extent allowed by law, the scope of the disclosure is to be
determined by the broadest permissible interpretation of the following
claims and their equivalents, and shall not be restricted or limited by
the foregoing detailed description.

Patent applications by Andrew M. Robb, Ravensdale, WA US

Patent applications by The Boeing Company

Patent applications in class Conductor or circuit manufacturing

Patent applications in all subclasses Conductor or circuit manufacturing